RAE 104 AIRFOIL (rae104-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: RAE 104 AIRFOIL (rae104-il) Reynolds number: 50,000 Max Cl/Cd: 24.61 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rae104-il-50000.txt Download as CSV file: xf-rae104-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: RAE 104 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.5552 0.10021 0.09318 0.0059 1.0000 0.3899
-8.750 -0.5339 0.09512 0.08804 0.0058 1.0000 0.3966
-8.500 -0.6580 0.07848 0.07189 -0.0203 1.0000 0.2113
-8.250 -0.7003 0.07079 0.06404 -0.0238 1.0000 0.1725
-8.000 -0.7252 0.06446 0.05729 -0.0236 1.0000 0.1470
-7.750 -0.7340 0.05947 0.05185 -0.0218 1.0000 0.1339
-7.500 -0.7310 0.05492 0.04703 -0.0199 1.0000 0.1275
-7.250 -0.7314 0.05097 0.04245 -0.0168 1.0000 0.1202
-7.000 -0.7226 0.04712 0.03829 -0.0145 1.0000 0.1170
-6.750 -0.7137 0.04362 0.03431 -0.0117 1.0000 0.1143
-6.500 -0.7014 0.04046 0.03068 -0.0091 1.0000 0.1137
-6.250 -0.6870 0.03793 0.02776 -0.0068 1.0000 0.1179
-6.000 -0.6712 0.03563 0.02486 -0.0042 1.0000 0.1218
-5.750 -0.6482 0.03274 0.02184 -0.0030 1.0000 0.1263
-5.500 -0.6247 0.03062 0.01940 -0.0015 1.0000 0.1348
-5.250 -0.5995 0.02862 0.01726 -0.0005 1.0000 0.1502
-5.000 -0.5657 0.02633 0.01519 -0.0005 1.0000 0.1740
-4.750 -0.5408 0.02433 0.01348 0.0007 1.0000 0.2186
-4.500 -0.5345 0.02172 0.01191 0.0045 1.0000 0.3119
-4.250 -0.2405 0.02815 0.01849 -0.0163 1.0000 0.9504
-4.000 -0.1377 0.02591 0.01561 -0.0323 1.0000 0.9949
-3.750 -0.1115 0.02489 0.01436 -0.0342 1.0000 1.0000
-3.500 -0.0978 0.02426 0.01363 -0.0334 1.0000 1.0000
-3.250 -0.0841 0.02371 0.01298 -0.0325 1.0000 1.0000
-3.000 -0.0706 0.02323 0.01242 -0.0314 1.0000 1.0000
-2.750 -0.0575 0.02281 0.01194 -0.0302 1.0000 1.0000
-2.500 -0.0451 0.02246 0.01153 -0.0288 1.0000 1.0000
-2.250 -0.0336 0.02216 0.01118 -0.0271 1.0000 1.0000
-2.000 -0.0234 0.02193 0.01093 -0.0252 1.0000 1.0000
-1.750 -0.0146 0.02176 0.01076 -0.0230 1.0000 1.0000
-1.500 -0.0078 0.02165 0.01065 -0.0205 1.0000 1.0000
-1.250 -0.0033 0.02160 0.01061 -0.0176 1.0000 1.0000
-1.000 -0.0010 0.02160 0.01061 -0.0143 1.0000 1.0000
-0.750 -0.0001 0.02162 0.01062 -0.0108 1.0000 1.0000
-0.500 0.0001 0.02165 0.01065 -0.0072 1.0000 1.0000
-0.250 0.0001 0.02167 0.01066 -0.0036 1.0000 1.0000
0.000 0.0000 0.02167 0.01067 0.0000 1.0000 1.0000
0.250 -0.0001 0.02167 0.01066 0.0036 1.0000 1.0000
0.500 -0.0001 0.02165 0.01065 0.0072 1.0000 1.0000
0.750 0.0001 0.02162 0.01062 0.0108 1.0000 1.0000
1.000 0.0010 0.02160 0.01061 0.0143 1.0000 1.0000
1.250 0.0033 0.02160 0.01061 0.0175 1.0000 1.0000
1.500 0.0077 0.02165 0.01065 0.0205 1.0000 1.0000
1.750 0.0146 0.02175 0.01075 0.0230 1.0000 1.0000
2.000 0.0234 0.02192 0.01093 0.0252 1.0000 1.0000
2.250 0.0336 0.02216 0.01118 0.0271 1.0000 1.0000
2.500 0.0451 0.02245 0.01153 0.0287 1.0000 1.0000
2.750 0.0575 0.02280 0.01193 0.0302 1.0000 1.0000
3.000 0.0707 0.02322 0.01241 0.0314 1.0000 1.0000
3.250 0.0842 0.02370 0.01297 0.0325 1.0000 1.0000
3.500 0.0979 0.02425 0.01361 0.0334 1.0000 1.0000
3.750 0.1117 0.02488 0.01435 0.0342 1.0000 1.0000
4.000 0.1368 0.02587 0.01556 0.0325 0.9953 1.0000
4.250 0.2402 0.02814 0.01848 0.0163 0.9507 1.0000
4.500 0.5345 0.02172 0.01191 -0.0045 0.3117 1.0000
4.750 0.5408 0.02433 0.01348 -0.0007 0.2187 1.0000
5.000 0.5658 0.02633 0.01518 0.0005 0.1744 1.0000
5.250 0.5992 0.02856 0.01723 0.0005 0.1507 1.0000
5.500 0.6248 0.03062 0.01936 0.0015 0.1349 1.0000
5.750 0.6482 0.03273 0.02184 0.0030 0.1264 1.0000
6.000 0.6714 0.03564 0.02486 0.0041 0.1219 1.0000
6.250 0.6869 0.03793 0.02776 0.0068 0.1178 1.0000
6.500 0.7014 0.04046 0.03068 0.0091 0.1137 1.0000
6.750 0.7137 0.04362 0.03430 0.0117 0.1143 1.0000
7.000 0.7225 0.04714 0.03831 0.0145 0.1170 1.0000
7.250 0.7315 0.05097 0.04245 0.0168 0.1202 1.0000
7.500 0.7311 0.05491 0.04702 0.0198 0.1276 1.0000
7.750 0.7344 0.05947 0.05184 0.0218 0.1340 1.0000
8.000 0.7252 0.06446 0.05729 0.0236 0.1471 1.0000
8.250 0.6376 0.05857 0.05215 0.0282 0.1556 1.0000
8.500 0.6119 0.06425 0.05793 0.0287 0.1677 1.0000
8.750 0.5772 0.07032 0.06400 0.0277 0.1809 1.0000
9.000 0.5466 0.09960 0.09252 -0.0059 0.3881 1.0000
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Polar data table (+)
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